Combustion engines are machines that convert chemical energy stored in fuel into mechanical energy useful for generating electricity, producing thrust, or otherwise doing work. These engines typically include several cooperative sections that contribute in some way to this energy conversion process. In gas turbine engines, air discharged from a compressor section and fuel introduced from a fuel supply are mixed together and burned in a combustion section. The products of combustion are harnessed and directed through a turbine section, where they expand and turn a central rotor.
A variety of combustor designs exist, with different designs being selected for suitability with a given engine and to achieve desired performance characteristics. One popular combustor design, known as a can-annular type design, comprises in each of a plurality of arranged “cans” a centralized pilot burner (hereinafter referred to as a pilot burner or simply pilot) and a number of main fuel/air mixing apparatuses. The main fuel/air mixing apparatuses are arranged circumferentially around the pilot burner. With this design, a central pilot flame zone and a mixing region are formed. During operation, the pilot burner selectively produces a stable flame in the pilot flame zone, while the fuel/air mixing apparatuses each produce a mixed stream of fuel and air in the above-referenced mixing region. The stream of mixed fuel and air flows out of the mixing region, past the pilot flame zone, and into a main combustion zone, where additional combustion occurs. Energy released during combustion is captured by the downstream components to produce electricity or otherwise do work.
In order to ensure optimum performance of a common combustor, it is generally preferable that the respective fuel-and-air streams are well mixed to avoid localized, fuel-rich regions. As a result, efforts have been made to produce combustors with essentially uniform distributions of fuel and air. Swirler elements, for example, are often used to produce a stream of fuel and air in which air and injected fuel are evenly mixed.
Gas turbine technology has evolved toward greater efficiency and also to accommodate environmental standards in various nations. One aspect in the evolution of designs and operating criteria is the use of leaner gas air mixtures to provide for increased efficiency and decreased emissions of NOx and carbon monoxide. Combustion of over-rich pockets of fuel and air leads to high-temperature combustion that produces high levels of unwanted NOx emissions.
Also, a key objective in design and operation of gas turbine combustors is the stability of the flame and, related to that, the prevention of flashbacks. A flashback occurs when flame travels upstream from the combustion zone in the combustion chamber and approaches, contacts, and/or attaches to, an upstream component. Although a stable but lean mixture is desired for fuel efficiency and for environmentally acceptable emissions, a flashback may occur at times more frequently with a lean mixture, and particularly during unstable operation. For instance, the flame in the combustion chamber may progress backwards and rest upon, for a period, a base plate which defines the upstream end of the combustion chamber. Less frequently, the flame may flash back into a fuel/air mixing apparatus, damaging components that mix the fuel with the air.
A multitude of factors and operating conditions provide for efficient and clean operation of the gas turbine combustor area during ongoing operation. Not only is the fuel/air mixture important. Also relevant to gas turbine operation are the shape of the combustion area, the arrangement of assemblies that provide fuel, and the length of the combustor that provides varying degrees of mixing. Given the efficiency and emissions criteria, the operation of gas turbines requires a balancing of design and operational approaches to maintain efficiency, meet emission standards, and avoid damage due to undesired flashback occurrences.
Also relevant to design and operation of gas turbine combustors is the avoidance of breakage of components, such as due to stress from vibration and cyclic stress, such as may come from having a first fundamental mode (i.e., a first natural frequency) within the range of the combustion dynamics.
The type of fuel/air mixing apparatus, and how it operates in relationship to other components, is one of the key factors in proper operation of current gas turbines. A common type of fuel/air mixing apparatus is known as a main swirler assembly. A main swirler assembly is comprised in part of a substantially hollow inner body that comprises stationary flow conditioning members (such as vanes) that create a turbulent flow. Fuel from a fuel nozzle is added before or into this turbulent air stream and mixes to a desired degree within a period of time and space so that it is well-mixed upon combustion in the downstream combustion chamber. Also, in typical arrangements, a main swirler assembly also is comprised of an outer downstream element known as an annulus casting. An annulus casting (referred to in some references as a “sleeve”) surrounds a downstream section of the inner body, forming a channel for air flow known as the flashback annulus. In a typical arrangement, a quantity, such as eight, swirler assemblies are arranged circumferentially around the central pilot burner. The pilot burner burns a relatively richer mixture than is provided by the radially arranged swirler assemblies. In a typical arrangement of a can-annular design of a gas turbine, 16 combustor cans, each having eight main swirler assemblies disposed around a central pilot burner, are arranged annularly and collectively provide combusted gases to a turbine.
As shown in FIG. 1, an example of a prior art gas turbine combustor 10 comprises a nozzle housing 12 having a nozzle housing base 14. A diffusion fuel pilot nozzle 16, having a pilot fuel injection port 18, extends through nozzle housing 12 and is attached to nozzle housing base 14. In the shown configuration, main fuel nozzles 20, each having at least one main fuel injection port 22, extend substantially parallel to pilot nozzle 16 through nozzle housing 12 and are attached to nozzle housing base 14. Fuel inlets 24 provide fuel 26 to main fuel nozzles 20. A main combustion zone 28 is formed within a liner 30 peripheral and downstream to a pilot flame zone 38. A pilot cone 32, having a diverged end 34, projects from the vicinity of pilot fuel injection port 18 of pilot nozzle 16. Diverged end 34 is downstream of main swirler assemblies 36. The pilot flame zone 38 is formed within pilot cone 32 adjacent to and upstream to main combustion zone 28.
Compressed air 40 from compressor 42 flows between support ribs 44 through main swirler assemblies 36. Each main swirler assembly 36 is substantially parallel to pilot nozzle 16 and adjacent to main combustion zone 28. Within each main swirler assembly 36, a plurality of swirler vanes 46 generate air turbulence upstream of main fuel injection ports 22 to mix compressed air 40 with fuel 26 to form a fuel/air mixture 48. Fuel/air mixture 48 is carried into main combustion zone 28 where it combusts. Compressed air 50 enters pilot flame zone 38 through a set of stationary turning vanes 52 located inside pilot swirler 54. Compressed air 50 mixes with pilot fuel 56 within pilot cone 32 and is carried into pilot flame zone 38 where it combusts.
FIG. 2 shows a detailed view of an exemplary prior art main swirler assembly 36. As shown in FIG. 2, main swirler assembly 36 is substantially cylindrical in shape, having a flared inlet end 58 and a tapered outlet end 60. A plurality of swirler vanes 46 are disposed circumferentially around the inner perimeter 62 of fuel swirler 36 proximate flared end 58. In the shown configuration, main swirler assembly 36 surrounds main fuel nozzle 20 proximate main fuel injection ports 22. Main swirler assembly 36 is positioned with swirler vanes 46 upstream of main fuel injection ports 22 and tapered end 60 adjacent to main combustion zone 28. Flared inlet end 58 is adapted to receive compressed air 40 and channel it into fuel swirler 36. Tapered outlet end 60 is adapted to fit into annulus casting 64 by weld connection, forming a channel known as a flashback annulus 65 through which air passes. Swirler vanes 46 are attached to a hub 66. Hub 66 surrounds main fuel nozzle 20.
FIG. 3 shows an upstream view of combustor 10. Pilot nozzle 16 is surrounded by pilot swirler 54. Pilot swirler 54 has a plurality of stationary turning vanes 52. Pilot nozzle 16 is surrounded by a plurality of main fuel nozzles 20. A main swirler assembly 36 surrounds each main fuel nozzle 20. Each main swirler assembly 36 has a plurality of swirler vanes 46. The diverged end 34 of pilot cone 32 forms an annulus 68 with liner 30. Main fuel swirlers 36 are upstream of diverged end 34. A fuel/air mixture flows through annulus 68 (out of the page) into main combustion zone 28 (not shown in FIG. 3).
As viewable in FIG. 2, main swirler assembly 36 is attached to liner 30 via attachments 70 and a conventional base plate 72. With respect to the latter manner of attachment, the distal end of annulus casting 64 is adjacent to the conventional base plate 72 as shown in FIG. 2. The conventional base plate 72 has a plurality of openings (such as 77 in FIG. 2) that are defined by downstream-oriented lips (such as 76 in FIG. 2). The distal end of annulus casting 64 and the downstream-oriented lip 76 of the conventional base plate 72 typically do not come into contact and are actually spaced up to approximately 10 mils (0.010 inches) apart in prior art embodiments. FIG. 3 shows a circular array of six swirlers, but other quantities, such as a series of eight swirlers, can be employed.
The other manner of attaching the main swirler assembly 36 to liner 30 is by way of attachments 70. In some prior art designs, attachments 70 comprised dual straight pins, each pin being welded at one end to liner 30 and at the other end to the main swirler assembly 36 as shown in FIG. 2. In other designs, hourglass-shaped pins (as shown in FIG. 4) provided larger weld areas on both the main swirler assembly 36 and the liner 30. Under certain operating conditions, both types of designs were observed to have fatigue-related failures. This led to consideration that the problem of such failures was related to the main swirler assemblies having a natural frequency within the range of the natural frequency of the combustion events in the combustor, and to the solution of this identified problem as disclosed herein.
More particularly, without being held to a particular theory, it is believed that some fatigue failures stem from a swirler's exposure to vibrational forces generated during combustor operation. Dominant combustion dynamics typically range from approximately 110-150 Hz, although variations outside this range are possible depending on the system design. The main swirler assembling in prior art swirler/base plate arrangements, when adjacent to the base plate, generally has a first fundamental mode (i.e., a natural frequency) of approximately 145 Hz, falling within the typical vibrational range of such combustion dynamics. Consequently, when a swirler is subjected to such forces, the swirler will resonate, and, it is believed, repeated resonance of the swirler ultimately fatigues the weld joints of the support pins.
Thus, high cycle fatigue failures are believed to be a recurring problem with respect to swirlers and other turbo machinery components. The problem has been exacerbated by combustion design changes to reduce emissions and increase efficiency. These design changes have increased the severity of the combustion dynamics, requiring more robust swirler assemblies. Therefore, there is a continuing need for a swirler assembly that can avoid vibration-induced resonance, that can further enhance the inherent damping characteristics of the main swirler to constrain any vibratory motion, and further that has a design that reduces or eliminates the occurrence of flashback.
Various approaches to reduce or eliminate flashback in modern gas turbine combustion systems have been attempted. Since the prevention or elimination of flashbacks is a multi-factorial issue and also relates to various aspects of the design and operation of the gas turbine combustion area, a range of approaches has been attempted. These approaches often inter-relate with one another.
One example of approaches to reach a balance among the needs to reduce flashbacks, maintain reasonable initial costs, maintain operating efficiency, and reduce downtime and costs due to component failure is provided in U.S. Pat. No. 6,705,087. This and all other patents, patent applications, patent publications, and other publications referenced herein are hereby incorporated by reference in this application in order to more fully describe the state of the art to which the present invention pertains, and to provide such teachings as are generally known to those skilled in the art.
The inventor of the present invention has appreciated the importance considering the durability criterion along with reduction of flashback. As a solution to the problem of balancing these criteria, a solution in the form of the present invention was attained. More particularly, the present invention provides a solution toward obtaining an operationally stable, flashback-resistant main fuel/air mixing apparatus, such as a swirler assembly, that is structurally stable and has an elevated natural frequency outside the range of the normal dominant combustion dynamics frequency.